1. Field of the Invention
The present invention relates generally to high temperature ceramic insulation materials applied to high strength ceramic substrates to form a hybrid structure designed for use in high temperature applications, especially gas turbines. More specifically, a hybrid ceramic structure is disclosed where the thermal insulating material is also thermally stable and erosion resistant and protects the underlying structural material from high temperatures in (for example) a turbine environment.
2. Background Information
Combustion turbines comprise a casing or cylinder for housing a compressor section, a combustion section and a turbine section. A supply of air is compressed in the compressor section and directed unto the combustion section. Fuel enters the combustion section by means of a nozzle. The compressed air enters the combustion inlet and is mixed with the fuel. The air/fuel mixture is then combusted to produce high temperature and high pressure gas. This working gas is then ejected past the combustor transition and injected into the turbine section of the turbine.
The turbine section comprises rows of vanes which direct the working gas to the airfoil portions of the turbine blades. The working gas slows through the turbine section, causing the turbine blades to rotate, thereby turning the rotor. The rotor is also attached to the compressor section thus turning the compressor and also an electrical generator for producing electricity.
A high efficiency of a combustion turbine is achieved by heating the gas flowing through the combustion section to as high a temperature as is feasible. The hot gas, however, heats the various metal turbine componentsxe2x80x94such as the combustor, transition ducts, vanes, ring segments and turbine bladesxe2x80x94that it passes when flowing through the turbine.
Accordingly, the ability to increase the combustion firing temperature is limited by the ability of the turbine components to withstand increased temperatures. Metallic structures within a turbine, whether with or without thermal barrier coatings (TBCs), require cooling. Thin layers of TBCs on the metallic structures are commonly used to protect critical components from premature breakdown due to increased temperatures to which the components are exposed. Generally, TBCs extend the life of critical components by reducing the rate of metal waste (through spalling) by oxidation and protecting underlying high strength structural superalloy substrates from intense heat.
Various cooling methods have been developed to cool turbine hot parts. These methods include open-loop air cooling techniques and closed-loop cooling systems. Both techniques, however, involve significant design complexity, have considerable installation and operating costs, and often carry attendant losses in turbine efficiency. For some applications, steam cooling is also being used which is more expensive and more complicated than air-cooling.
Conventional state-of-the-art first row turbine vanes are fabricated from single-crystal superalloy castings with intricate cooling passages and with external TBCs. Not only are these components expensive to manufacture, but with ever-increasing gas path temperatures, their ability to be effectively cooled is limited. These vanes are subjected to high velocity, high temperature gases under high-pressure conditions.
The TBC coating thickness on the turbine vanes and blades must be limited to prevent residual stress buildup and spallation. Potential coating compositions are generally limited to high expansion materials to minimize thermal expansion mismatch between the TBC and substrate metal. Also, the TBC coating has limited durability due to high thermally induced stresses caused by both the thermal expansion mismatch and metal substrate oxidation.
Currently the state of art TBC technology also is limited to surface temperatures of less than 1200xc2x0 C. for long term use. Also, current TBC compositions are limited to high coefficient of thermal expansion materials, such as ZrO2, to minimize the thermal expansion mismatch between the superalloy and the TBC; at temperatures less than 1200xc2x0 C., these TBCs can sinter to near theoretical density, which can lead to spallation. As stated above active cooling of the components is required.
In Advanced Turbine systems (ATSs), the temperature demands of operation and the limits of ATS state-of-the-art materials, may lead to eventual failure of even the most sophisticated high temperature TBCs. This, in turn, can result in premature failure of the critical components and therefore, potential failure of the turbine, interruption in the power supply and expensive repair costs. It is, therefore, desirable to provide turbine components that can (1) withstand high temperatures without the use of the thermal barrier coatings and (2) which substantially reduce the need for cooling.
Other materials for thermal insulation are fibrous ceramic insulating materials. A major drawback of these materials, however, is that they have low densities which lead to very poor erosion resistance. Therefore, fibrous ceramic insulating materials are inapplicable to high velocity gas flow applications.
Monolithic tiles are another material that could possibly be used for protecting critical components in high temperature conditions. These tiles have good erosion resistance and insulating properties, however, they are susceptible to thermal shock damage and catastrophic failure. It is, therefore, desirable to provide insulating materials that can withstand high temperatures without the use of thermal barrier coatings, fibrous ceramic insulating materials, or monolithic ceramic tiles.
Commercially available ceramic matrix composites (CMCs), for example, were thought to have some potential applications in gas turbines, but they are limited in their exposure to temperatures near 1200xc2x0 C. for long periods of time, that is, greater than 10,000 hours for gas turbines as needed for power generation. In addition, CMCs cannot be effectively cooled under high temperature conditions (greater than 1400xc2x0 C.) or high heat flux conditions due to their relatively low thermal conductivity and inability to fabricate intricate cooling passages.
What is needed is a structure to replace prior art TBC coated metal substrates for hot gas path components in turbine engines. Therefore, it is an object of this invention to provide a material structure that can significantly reduce component cooling requirements, that as compared to the prior/current technology can provide a high temperature erosion resistant material, and that can withstand high temperature environments without degradation.
These and other objects of the invention are accomplished by providing a thermally stable engineered layered ceramic structure, henceforth known as the xe2x80x9cHybrid Ceramicxe2x80x9d that operates with two aspects. One being a high temperature resistant insulating layer attached to a second more rigid structural layer. The insulating layer is temperature stable (i.e., microstructurally stable and effectively non-sintering), thermally insulating, low elastic modulus ceramic. The structural layer has a lower temperature stability compared to the insulating layer but is mechanically load bearing with a higher elastic modulus than the insulating layer. The proposed system functions similarly to a conventional TBC coated superalloy system but has many more advantages.
The hybrid ceramic is designed to operate under high heat flux conditions with the insulating layer exposed to high temperature gases or other fluid media and with cooling applied to the structural member through cooling fluid means. Thus the system operates under a thermal gradient with the insulating layer having a significantly higher temperature than the cooled structural member. The specific design of the hybrid system is such that the structural member is maintained at a sufficiently low temperature where its mechanical properties are adequate for the load bearing requirements of the application and its microstructural stability is maintained for the desired lifetime of the component.
The hybrid ceramic system of the present invention system is of a compatible ceramic composition. Thus the thermo-mechanical mismatch between the structural layer and the insulating layer is minimized, meaning that the insulating layer in the hybrid ceramic can be much thicker than the insulating ceramic layer of typical TBC/metal structures. Thus, much greater thermal protection is provided to the substrate material, allowing the use of lower temperature capable structural materials in the same high temperature environment (for example, using a 1200xc2x0 C. capable CMC in a  greater than 1600xc2x0 C. environment).
Another feature of the present invention is that the insulating layer is not as limited in material selection and capability as that for conventional metal/TBC systems and can, thus, be comprised of a material with much higher temperature stability. This capability means that the present invention provides the capability to withstand much higher temperatures than conventional metal/TBC systems can withstand. The thermal stability of the insulating layer is a key feature of the invention, minimizing stresses resulting from sintering shrinkage strains and maintaining the integrity of the insulating layer and thus the integrity of the hybrid ceramic structure over an extended operating life.
A further feature of the present invention is that the structural layer material is comprised of a ceramic rather than a metal so that it can also impart improved thermal properties, in the form of increased thermal resistance. This capability, which allows the use of low thermal conductivity structural layers such as oxidexe2x80x94oxide CMC materials, reduces the heat withdrawal from the engine system, thereby reducing cooling air needs and increasing the power output and thermal efficiency of the engine.
Yet another feature is that the insulating layer material can be selected to be preferentially abradable so that the hybrid system can be use as an abradable sealing component for the ends of the blades.
A preferred embodiment of the invention consists of an underlying structural layer and a protective thermal insulating layer. The structural layer is made of a continuous fiber oxidexe2x80x94oxide ceramic matrix composite that is micro-structurally stable and possesses long term mechanical strength and durability up to about 1200xc2x0 C. This layer is of the order of 3 to 10 mm thick or can be thicker depending upon the application.
The thermal insulating layer is comprised of closely packed thermally stabilized (to 1700xc2x0 C.) ceramic oxide spheres. This layer is of the order of 2 to 5 mm thick or can be thicker depending upon the application. Also, the insulating layer can be comprised of hollow or partially hollow (including porous core) sphere-based structures, the walls of which are sufficiently thin to impart excellent abradability to the system. This hybrid structure of the present invention has the inherent advantage that it can withstand exposure to hot gas temperatures close to 1700xc2x0 C. (i.e. greatly in excess of conventional systems). It can be engineered by controlling the relative thickness of the structural layer and the insulating layer so that the thermal protection afforded to the structural layer is of several hundred centigrade degrees (of the order of 200 to 700 centigrade degrees for high heat flux turbine applications). The structural material, therefore, operates well within the regime of its long term mechanical capabilities.
Although the optimum properties are provided by this specific combination of material, specifically required subsets of these properties can be generated using other coatings and substrates. The invention can employ alternative substrate materials and alternative coatings to yield similarly functioning thermo-mechanical ceramic hybrid systems.
This invention provides hybrid ceramic structure that enables the use of a ceramic composite in application environments, such as gas turbines, where normal materials (including monolithic ceramics or stand-alone CMCs) could not be used. The hybrid ceramic uses the structure of two or more ceramic materials bonded/attached together to present the insulating material to the hot gas environment and the structural material to the colder (cooling medium) environment. This hybrid ceramic exposes the special insulating material to temperatures that cannot be withstood by existing structural ceramic materials, such as ceramic matrix composites or monolithic ceramics. It can significantly reduce component cooling requirements, up to about 90% for gas turbine hot gas path components, as compared to the prior/current technology. The insulating layer can also be engineered to provide a high temperature erosion resistant abradable system, which can withstand high temperature environments that degrade the prior/current technology.
This invention allows (but is not limited to) use of structural layers which are not suitable at temperatures over 1200xc2x0 C. even though the system is exposed to a 1600xc2x0 C.-1700xc2x0 C. environment. Thus, common relatively inexpensive materials such as ceramic matrix composites (CMCs), fibrous ceramics and monolithic ceramic can be utilized as the structural layer, when operating in a turbine environment where the insulating layer is exposed to temperatures from 1400xc2x0 C. to 1700xc2x0 C. Preferably, the insulating layer is more than 20% porous, and the structural layer is less than 20% porous. The invention can be applied to several gas turbine components of several types (such as blade and vane airfoils, vane platforms, combustors, ring segments or transitions), as well as a variety of applications wherein high temperature, high hot gas velocities, and/or high heat fluxes are required.